Blade component

ABSTRACT

A blade component has a longitudinal axis and extends between a root end and a tip end. The component comprises a composite laminate structure, the laminate structure comprising a plurality of plies of fibres in a matrix, wherein all the plies are arranged such that the fibres in respective plies are oriented symmetrically relative to the component axis at respective layup angles, the layup angles being in the range of 19° to 25° and −19° to −25° respectively relative to the component axis.

FOREIGN PRIORITY

This application claims priority to European Patent Application No.16306365.4 filed Oct. 17, 2016, the entire contents of which isincorporated herein by reference.

TECHNICAL FIELD

The present disclosure relates to blade components such as propellerblade structural spars.

BACKGROUND

Blade components such as propeller blade structural spars may include alaminate structure to provide improved mechanical characteristics of thecomponent. The laminate structure may be formed from several layers orplies, each ply comprising fibres aligned in the same direction acrosstheir respective surfaces.

In order to accommodate the various loads in the structure, theorientation of the fibres typically differs in the plies. In a knownlay-up, as shown in FIG. 1 a first ply 12 has fibres aligned with theaxis 11 of the component, i.e. a 0° orientation and second and thirdplies 14, 16 oriented at +/−45° from the axis 11 of the component.

In such an arrangement, centrifugal loads and aero bending momentresistance and stiffness are sustained by the 0° fibres of the first ply12, while the torsional loading and torsional stiffness are sustained bythe +/−45° fibres of the second and third pliers 14, 16.

However, in order to provide optimum performance the +/−45° plies 14, 16and 0° pliers 12 should be mixed and equally distributed along thelay-up thickness to avoid interlaminar shear overstresses. The lay-upshould also be arranged symmetrically about the axis to avoid thermaloverstresses and material distortion of the component. However, duringmanufacture, equal distribution and exact symmetry can be difficult toattain. Furthermore, when braiding the layers, the 0° and +/−45° layersrequire two different braiding machines. Each machine requires differentconfigurations to manufacture these two different angles.

SUMMARY

In accordance with the disclosure, there is provided a blade componenthaving a longitudinal axis and extending between a root end and a tipend, the component comprising a composite laminate structure. Thelaminate structure comprises a plurality of plies of fibres in a matrix,wherein all the plies are arranged such that the fibres in respectiveplies are oriented symmetrically relative to the component axis atrespective layup angles. The layup angles are in the range of 19° to 25°and −19° to −25° respectively relative to the component axis.

The plurality of plies may comprise one or more first plies eachcomprising fibres aligned at a layup angle α relative to the componentaxis and one or more second plies each comprising fibres aligned at alayup angle of −α relative to the component axis.

The plurality of plies may further comprise one or more third pliescomprising fibres aligned at a layup angle of β relative to thecomponent axis and one or more fourth plies comprising fibres aligned ata layup angle of −β relative to the component axis and wherein β isdifferent from α and wherein α and β are both in the range of 19° to 25°and −19° to −25°.

Alternating plies may be arranged symmetrically relative to thecomponent axis.

The layup angle may be in the range ±20° to 23°.

There is also provided a method of manufacturing a blade componenthaving a longitudinal axis and extending between a root end and a tipend. The method comprises laying a plurality of plies of fibres on acore, all the plies being arranged such that the fibres in the plies areoriented symmetrically relative to the component axis at respectivelayup angles, the layup angles being in the range of 19° to 25° and −19°to −25° respectively relative to the component axis. The method furthercomprises curing the component with the plies to form a laminatedstructure on the core.

The plurality of plies may comprise one or more first plies eachcomprising fibres aligned at a layup angle α relative to the componentaxis and one or more second plies each comprising fibres aligned at alayup angle of −α relative to the component axis.

The plurality of plies may further comprise one or more third pliescomprising fibres aligned at a layup angle of β relative to thecomponent axis and one or more fourth plies comprising fibres aligned ata layup angle of −β relative to the component axis.

Alternating plies may be oriented symmetrically relative to thecomponent axis.

The layup angle may be in the range ±20° to 23°

The plies may be formed from sheets or tapes of fibre material.

The plies may be pre-impregnated with a resin or wherein a resin isapplied to the plies for curing.

The blade component may be a propeller blade component or a fan bladecomponent. For example, the blade component may be a blade spar. Thelaminate structure may be provided on a core of the spar.

BRIEF DESCRIPTION OF DRAWINGS

Some embodiments of the disclosure will now be described by way ofexample only and with reference to the accompanying drawings in which:

FIG. 1 shows a lay-up of fibre plies for a propeller blade componentaccording to the prior art;

FIG. 2 shows a lay-up of fibre plies for a propeller blade componentaccording to this disclosure;

FIG. 3 shows a propeller blade component and plies oriented at differentangles relative to the component axis; and

FIG. 4 shows a propeller blade component and composite tapes oriented atdifferent angles relative to the component axis.

DETAILED DESCRIPTION

With reference to FIGS. 2 to 4, FIG. 2 shows an exemplary laminatestructure 20 for a propeller blade component such as a structural spar30 as illustrated in FIGS. 3 and 4.

With reference to FIG. 3, the structural spar 30 includes a root 32 anda tip 34 and a body 33 extending from the root 32 to the tip 34. Thebody 33 extends along an axis 100 defining 0° orientation with respectto the ply lay-up. The axis may be in the perpendicular plane of thepropeller blade rotation axis or, in the case of curved blades, may havea local blade axis at any given point defined perpendicular to the localblade section. The root 32 comprises or is attached to a retentionelement 35 which retains the blade in a hub in use. The blade is usuallyrotatable about the axis 100, for example to allow the blade to befeathered.

Returning to FIG. 2, the laminate structure 20 includes a first ply 22and a second ply 24. The first ply 22 comprises a plurality of fibresshown schematically at 26. The fibres 26 may, for example, comprisecarbon fibres, glass fibres, aramid fibres or the like. The fibres 26 ofthe first ply 22 are aligned along a single direction shown by the arrow23. That is, a majority of or substantially all the fibres 26 arealigned in the same direction throughout the first ply 22. For example,at least 90% of the fibres are aligned in the same direction and up to10% of the fibres might be arranged at about 90° to the remainingfibres. The first ply 22 is oriented relative to a central axis 21 suchthat a layup angle +α is defined between the central axis 21 and thedirection 23 of the fibres 26. The axis 21 is aligned with alongitudinal axis of the component, for example the axis 100 of the spar30 as illustrated in FIGS. 3 and 4.

The second ply 24 comprised a plurality of fibres shown schematically at28. The fibres 28 may also, for example, comprise carbon fibres, glassfibres, aramid fibres. In the embodiment, the fibres 28 of the secondply 24 are substantially the same as the fibres 26 of the first ply 22.The fibres 28 of the second ply 24 are aligned along a single directionshown by the arrow 25. That is a majority of or substantially all thefibres 28 are aligned in the same direction throughout the second ply24. The second ply 24 is oriented relative to the central axis 21 suchthat a layup angle −α (i.e. the opposite of angle α) is defined betweenthe central axis 21 and the direction 25 of the fibres 28.

Thus the first ply 22 and the second ply 24 are arranged symmetricallyrelative to the central axis 21 of the structure.

The plies may be applied to a core to form the laminated structure. Inthe case of a spar as illustrated in FIG. 3, the pliers 22, 24 may beapplied to a lightweight core, for example a foam core. The first andsecond pliers 22, 24 may then be braided on to or otherwise attached tothe core. In embodiments, the plies 22, 24 might include dry fibres,such as carbon fibres, braided onto the outer surface of the core. Resinmay be subsequently injected into the pliers 22, 24. Alternatively, thepliers 22, 24 may be formed from a pre-impregnated fabric material, suchas a resin impregnated carbon fibre fabric.

As shown in FIGS. 3 and 4, the pliers 22, 24 might be attached to thecore in the form of sheets 38 or tapes 138. The pliers 22, 24 may beattached to the core such that they surround the entire circumference ofthe core.

In embodiments, the laminate structure 20 has a uniform thickness for agiven cross section along the spar 30. The first and second pliers 22,24 may also have the same thickness. It will be appreciated, however,that the thickness of the laminate structure 20 might be varied achievedby applying plies of uniform thickness only in specific areas on thecore, for example.

After resin injection or application of pre-impregnated material, thespar 30 is heated or cured to set the laminate structure 20.

Although the laminate structure 20 illustrated includes two pliers 22,24, it is envisaged that any number of plies may be used. For example,in a propeller blade spar the laminate structure 20 may include between15 and 30 plies. In another example, a laminate structure for use in afan blade may include up to and in excess of 80 plies.

The plies are arranged such that the fibres of all the plies areoriented at an angle of between 19° and 25° from the axis of the spar 30in either direction. In embodiments, the fibres may be oriented at anangle of 20°, 22° or 23° from the central axis 21. None of the plies isarranged such that the fibres are aligned with the axis 21 (i.e. at 0°).

In particular embodiments, there may only be first and second plies 22,24 having fibres 26, 28 arranged at the same angle relative to thecentral axis at respective layup angles of ±α, to the central axis 21.In embodiments, there may be multiple first and second plies with suchorientations.

In some multiple ply arrangements, the orientation of the plies relativeto the central axis 21 may alternate (for example, α−αα−α). In othermultiple ply arrangements, the order may not alternate between adjacentplies (for example αα−α−α)

Embodiments having pliers 22, 24 that are oriented at just one anglearound the axis 21 may be easier to form. In particular it may be easierto maintain the symmetry of the lay-up throughout the pliers 22, 24,particularly when compared to structures containing 0° fibres.

In yet further embodiments however, the laminate structure 20 mightinclude one or more first pliers 22 oriented at an angle α, one or moresecond pliers 24 oriented at −α, one or more third plies (not shown)oriented at an angle β and one or more fourth plies (not shown) orientedat −β, β being different from α.

Having the fibres 26, 28 of all the pliers 22, 24 oriented at an angleof between 19° and 25° from the axis 21 may improve the inter-laminarshear strength of the laminate structure 20 as the maximum angle betweentwo fibres of the structure is less than 90°.

Moreover, the curing thermal stressing between plies may be reduced oreven eliminated in the case of interlacing fibres.

A further advantage of embodiments having the plies oriented at anglesbetween 19 and 25° is that cutting of pliers 22, 24 during applicationto a core, will result in lower scrap material when compares to pliesthat are oriented at 45°, for example. This is best shown in FIG. 3 inwhich a first ply 38 is shown oriented at 19-25° from the blade axis anda second ply 36, outside the scope of the disclosure, is oriented at 45°from the blade axis. As can be seen from the lay-up of both plies, thefirst ply 38 requires less cutting of the ply material and thus lesswaste may be produced.

An embodiments where the plies are formed from tapes, is illustrated inFIG. 4. As shown, a first tape 136 is arranged at 45° relative to theaxis 100 and a second tape 138 is arranged at between 19 and 25°relative to the axis 100. The tapes may be automatically laid on a coresurface using a tape-laying apparatus. By having the angle between thetape 138 and the central axis 100 reduced, for example when the fibresof the tape are aligned along the length of the tape, fibre depositspeed can be increased and the steering of the apparatus across thesurface of the core may be simplified and reduced. Moreover, it will beseen that the second tape 136 extends along a greater length of the corestructure and therefore fewer tapes and less steering of the depositionapparatus may be required than in laying down tapes 136 at a greaterangle, outside the scope of this disclosure.

In embodiments where the plies are braided, only one braiding machinemay be needed to lay the plies, particularly in embodiments where thefibres are aligned symmetrically about the spar axis 100.

From the above, it will be recognised that there is proposed a laminatestructure wherein plies have fibres oriented at opposite angles relativeto a central axis, the angle being in the range of 19° to 25°. Inparticular embodiments, the fibres within the plies are oriented at asingle angle either side of the axis of the component to form asymmetrical laminate structure. The mechanical characteristics of suchlaminate structures have to been found to be comparable to the known0°±45° lay-up and further have considerable performance andmanufacturing advantages as described above. Table 1 shows a number ofmechanical characteristics of components having laminate structures inaccordance with the disclosure compared to the characteristics ofcomponents having the conventional 0°±45° lay-up. The characteristics ofthe components according to this disclosure were found to be comparablewithin acceptable limits for a variety of applications.

TABLE 1 [0°, +/−45°] [0°, +/−45°] [0°, +/−45°] [0°, +/−45°] LaminationType (50%/50%) [+/−23°] (50%/50%) [+/−22°] (60%/40%) [+/−20°] (50%/50%)[+/−20°] Type of Fibre Carbon HR Carbon HR Carbon IM Carbon IM Carbon HRCarbon HR Carbon IM Carbon IM Volume of Fibre 60% 60% 60% 60% 60% 60%60% 60% 1. Rigidity E1 (blade axis) 83 80 96 94 97 95 112 107 E2 (chorddirection) 24 10 26 9.4 22 10 23 10 G12 (torsional stiffness) 21 22 2424 18 19 21 21 Poisson's Ratio 0.7 1.5 0.75 1.7 0.68 1.4 0.72 1.6 2.Strength (Static Force) Ultimate blade axis (MPa) 680 490 1473 950 820630 1770 1140 First Ply Failure blade axis (Mpa) 470 490 770 950 550 630895 1140 Ultimate chord direction (Mpa) 290 47 379 47 240 46 340 46First Ply Failure chord direction 110 47 131 47 100 46 114 46 (Mpa) 3.Coefficient of Expansion Blade Axis −1.7 −3.5 −0.23 −2.15 −0.7 −2.9−0.25 −1.84 Chord Axis 23.5 22.7 5.23 16.4 9.8 24.2 6.3 17.1

While the disclosure has been particularly directed to a structural sparof a propeller blade, it may be used for other blade components, forexample an external skin for a propeller blade. The principles of theinvention may also be applied to fan blades, particularly thosemanufactures by Automatic Fibre Placement (AFP) process.

1. A blade component having a longitudinal axis and extending between aroot end and a tip end, the component comprising a composite laminatestructure, the laminate structure comprising a plurality of plies offibres in a matrix, wherein all the plies are arranged such that thefibres in respective plies are oriented symmetrically relative to thecomponent axis at respective layup angles, the layup angles being in therange of 19° to 25° and −19° to −25° respectively relative to thecomponent axis.
 2. A blade component as claimed in claim 1, wherein theplurality of plies comprises one or more first plies each comprisingfibres aligned at a layup angle α relative to the component axis and oneor more second plies each comprising fibres aligned at a layup angle of−α relative to the component axis.
 3. A blade component as claimed inclaim 2 wherein the plurality of plies further comprises one or morethird plies comprising fibres aligned at a layup angle of β relative tothe component axis and one or more fourth plies comprising fibresaligned at a layup angle of −β relative to the component axis andwherein β is different from α, wherein α and β are both in the range of19° to 25° and −19° to −25°
 4. A blade component as claimed in claim 1,wherein alternating plies are arranged symmetrically relative to thecomponent axis.
 5. A blade component as claimed in claim 1, wherein thelayup angle is in the range ±20° to 23°.
 6. A method of manufacturing ablade component having a longitudinal axis and extending between a rootend and a tip end, the method comprising: laying a plurality of plies offibres on a core, all the plies being arranged such that the fibres inthe plies are oriented symmetrically relative to the component axis atrespective layup angles, the layup angles being in the range of 19° to25° and −19° to −25° respectively relative to the component axis; andcuring the component with the plies to form a laminated structure on thecore.
 7. A method of manufacturing a blade component as claimed in claim6, wherein the plurality of plies comprises one or more first plies eachcomprising fibres aligned at a layup angle a relative to the componentaxis and one or more second plies each comprising fibres aligned at alayup angle of −α relative to the component axis.
 8. A method ofmanufacturing a blade component as claimed in claim 7, wherein theplurality of plies further comprises one or more third plies comprisingfibres aligned at a layup angle of β relative to the component axis andone or more fourth plies comprising fibres aligned at a layup angle of−β relative to the component axis.
 9. A method of manufacturing a bladecomponent as claimed in claim 6, wherein alternating plies are orientedsymmetrically relative to the component axis.
 10. A method ofmanufacturing a blade component as claimed in claim 6, wherein the layupangle is in the range ±20° to 23°
 11. A blade component or method asclaimed in claim 6, wherein the plies are formed from sheets or tapes offibre material.
 12. A method as claimed in claim 11, wherein the pliesare pre-impregnated with a resin or wherein a resin is applied to theplies for curing.
 13. A method of manufacturing a blade component asclaimed in claim 6, wherein the blade component is a propeller bladecomponent or a fan blade component.
 14. A method of manufacturing ablade component as claimed in claim 6, wherein the blade component is ablade spar.
 15. A blade component as claimed in claim 1, wherein theblade component is a propeller blade component or a fan blade component.16. A blade component as claimed in claim 1, wherein the blade componentis a blade spar.
 17. A blade component as claimed in claim 14, whereinthe laminate structure is provided on a core of the spar.